System and method for gyro enhanced vertical flight information

ABSTRACT

A system and associated method where a vertical flight dynamic sensor is stabilized using pitch rate gyro information combined with azimuth rate gyro information corrected by scaling the azimuth rate gyro information in accordance with feedback derived from the difference between the vertical flight dynamic sensor signal and the corrected vertical flight dynamic output. In one embodiment, the vertical flighy dynamic sensor may be a vertical speed sensor. In another embodiment, the vertical flight dynamic sensor may be a pitch attitude sensor derived from the difference between an inertial based acceleration signal and a velocity based acceleration signal. The velocity based acceleration signal may be based on airspeed or GPS velocity.

BACKGROUND

1. Field of the Invention

The present invention pertains to the field of aircraft flight controland instrumentation, more particularly to the field of vertical, orpitch axis control and instrumentation.

2. Background of the Invention

Aircraft flight control is a complex activity involving the managementof information from numerous instruments that keep track of aircraftflight states including three axes of position and velocity, three axesof aircraft attitude, wind, engine and airframe configuration, flaps,trim. For proper, efficient, and safe flight, the pilot must keep trackof and manage all of this information, often in a rapidly changingenvironment.

A pilot typically manages the vertical axis of an aircraft by referenceto a vertical gyro as the primary vertical axis instrument. The verticalgyro presents pitch attitude, which the pilot observes in order to steerthe elevator of the aircraft. To maintain level flight and holdaltitude, the pilot iteratively adjusts aircraft pitch attitude to avalue that is expected to yield the desired flight path and thenmonitors the altimeter and vertical speed indicator, airspeed indicatorand other instruments. As errors creep in, the pilot adjusts theaircraft attitude to fly back to the desired path. Thus, the pilot mustconstantly monitor the vertical gyro, altimeter, airspeed, and ifavailable, the vertical speed indicator, scanning frequently among themto maintain the desired flight path. Further, the pilot must maintainthe pitch axis while maintaining the roll axis, maintaining navigationupdates, observing engine instruments, and keeping in radio contact withflight control. Thus, maintaining the vertical axis potentiallycontributes significantly to the pilots work load and fatigue.

Reducing pilot fatigue and increasing air safety, potentially comes at acost. Precision aircraft instruments can be expensive to acquire,install, and maintain. Providing flight instruments, even fornon-instrument VFR (visual flight rules) flying can improve flightsafety by having instruments available for the occasional unanticipatedand unavoidable need for instrument flying. Further, air safety can beimproved by including redundant instruments, enabling an alternatesource of critical information in the case of failure of a particularinstrument. Adding redundant instruments, or even equipping an aircraftfor instrument flying can be expensive. Thus, techniques that enable theproduction of low cost flight instruments can materially improve flightsafety by encouraging the wider installation and use of flightinstruments.

Therefore, there is a need for flight instruments that provide efficientdisplay of flight information to reduce pilot work load and fatigue.Further, there is a need for flight instruments based on low costsensors to enable higher quality flight control at lower cost, improvingoverall flight safety.

BRIEF DESCRIPTION OF THE INVENTION

Briefly, the present invention is a system and associated method where avertical flight dynamic sensor is stabilized using pitch rate gyroinformation combined with azimuth rate gyro information corrected byscaling the azimuth rate gyro information in accordance with feedbackderived from the difference between the vertical flight dynamic sensorsignal and the corrected vertical flight dynamic output. In oneembodiment, the vertical flight dynamic sensor may be a vertical speedsensor. In another embodiment, the vertical flight dynamic sensor may bea pitch attitude sensor derived from the difference between an inertialbased acceleration signal and a velocity based acceleration signal. Thevelocity based acceleration signal may be based on airspeed or GPSvelocity.

BRIEF DESCRIPTION OF THE FIGURES

The present invention is described with reference to the accompanyingdrawings. In the drawings, like reference numbers indicate identical orfunctionally similar elements. Additionally, the left most digit(s) of areference number identifies the drawing in which the reference numberfirst appears.

FIG. 1 is an exemplary block diagram of a system providing gyro enhancedvertical information in accordance with the present invention.

FIG. 2A–FIG. 2E illustrate the filter response to various exemplaryvertical speed and pitch rate gyro signals.

FIG. 3 shows one embodiment of a vertical flight information sensor asused by the present invention.

FIG. 4 is a schematic diagram for an exemplary analog embodiment of thepresent invention.

FIG. 5 is an exemplary block diagram of an alternate verticalinformation source.

FIG. 6 illustrates a processor based embodiment of the presentinvention.

FIG. 7 illustrates an exemplary embodiment that includes velocityinformation.

FIG. 8 illustrates an exemplary embodiment of an attitude directionindicator (ADI) instrument in accordance with the present invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

The invention is a system and method for generating gyro enhancedvertical information (also called gyro stabilized vertical speed.) Inone embodiment, the vertical information is vertical speed information.In another embodiment, the vertical information is pitch attitudeinformation.

The combination of gyro based pitch attitude information and verticalspeed information in a single signal in accordance with the presentinvention results in a display that simplifies the task of flying anaircraft and reduces pilot work load and fatigue.

The invention presents a stabilized, flyable vertical speed display,free of the lag that is present in a pure vertical speed display.Without gyro stabilization, a vertical speed display includes too muchlag to be used to directly control an aircraft. The lag results from atleast two sources: the mass of the aircraft must be accelerated to a newvertical speed in response to an attitude control input, and thevertical speed sensor itself may have a lag in its response. If thepilot uses lagged vertical speed alone to control the aircraft, thepilot will typically over control, alternately pitching up too much andovercorrecting by pitching down too much, resulting in a hunting oroscillating maneuver that is at least uncomfortable, and potentiallydangerous. But, by using gyro stabilized vertical speed, the pilot mayfly directly according to the indicator without over controlling,resulting in smooth, stable flight.

Flying according to gyro enhanced vertical speed removes one layer ofmental activity involved in flying the aircraft. Using conventionalinstruments, the pilot adjusts pitch attitude repetitively withreference to a vertical gyro and a vertical speed indicator in order toachieve a desired vertical speed. Using a gyro enhanced vertical speedindicator, the pilot simply adjusts the pitch attitude of the aircraftuntil the display shows the desired vertical speed. Thus the use of gyroenhanced vertical speed reduces the scanning effort, reduces the numberof instruments that need to be scanned, and reduces the frequency ofscanning, reducing pilot workload and reducing pilot fatigue.

An instrument displaying gyro enhanced vertical speed can potentiallyreplace the vertical gyro as the primary vertical flight instrument,leaving the vertical gyro for backup or cross check use.

Gyro enhanced vertical speed may be produced using economical sensors.In accordance with one embodiment of the invention, two strap down rategyros and one economical pressure altitude sensor are provided. One ofthe rate gyros may also be used for a turn and bank, so that when turnand bank is provided only one additional gyro, the pitch rate gyro isprovided. The pressure altitude sensor may be economical because thealtitude sensor does not have to be precisely calibrated for altitudedisplay.

In accordance with the invention, a feedback signal based on thedifference between the vertical speed sensor and the displayed verticalspeed is used to scale an azimuth rate signal to correct the pitch rategyro signal in turns. The feedback correction enables low cost designwith minimum number of sensors for more reliable operation.

The design includes inherent safety properties because the display isbased on instantaneous sensors and can recover rapidly from out of rangeoperation. A pilot should not intentionally exceed 45 degrees roll, butthe instrument can recover in a few seconds even from extreme invertedattitude. Some conventional gyros may hit limit stops that causedeflections requiring several minutes to recover.

The potential low cost of an instrument in accordance with the presentinvention improves flight safety by enabling instrument flight in moreaircraft and enables redundant instruments in more aircraft. Flightsafety is further improved through a reduction in pilot workload andfatigue.

FIG. 1 is an exemplary block diagram of a system providing gyro enhancedvertical information in accordance with the present invention. Theinvention will first be described wherein the vertical information isvertical speed information as from a pressure altitude based verticalspeed sensor. Referring to FIG. 1, the vertical speed sensor 102produces a vertical speed signal 104, which is summed with anenhancement signal 110 from a pitch rate gyro 108 and a correctionsignal 128. The summed result is processed by a filter 112 to produce agyro enhanced vertical speed signal 114 (also called filtered signal.)The gyro enhanced vertical speed signal 114 is used to control a display116, where the signal 114 may be observed by the pilot and used tocontrol an aircraft. Alternatively, the signal 114 may be included in anautopilot system for control of the aircraft (not shown.)

Vertical speed information 102, by itself, is difficult to use to fly anaircraft because vertical speed information 102 lags significantlybehind the pitch control action used to control the vertical axis. Bycombining pitch rate gyro information 110 with vertical speedinformation 104 using a complementary filter, the fast responding pitchrate gyro information 110 may be used to place the aircraft in anattitude that will result in the desired vertical speed. The longer termvertical speed information 104 will then indicate the achieved verticalspeed.

For straight and level flying, vertical speed 104 and pitch rate gyroinformation 110 are sufficient; however, in a turn, the pitch rate gyro108 is also responsive to the turn rate. In fact, the pitch rate gyro108 is responsive to the azimuth rate multiplied by the sine of the bankangle. The azimuth rate response is undesired in the pitch axis display.If not corrected, the azimuth rate component of the pitch rate gyrosignal 110 would add to the displayed vertical speed, producing anexcessively high vertical speed indication, causing the pilot to pitchdown and lose altitude.

Thus, a correction signal 128 is subtracted from the pitch rate gyrosignal 110 to remove the azimuth rate response component of the pitchrate gyro signal 110. The correction signal 128 is produced by scalingthe absolute value 124 of an azimuth rate gyro signal 123 by adifference signal 120. The difference signal 120 is the differencebetween the displayed gyro enhanced vertical speed signal 114 (filteredsignal 114) and the vertical speed signal 104 from the vertical speedsensor 102.

The azimuth rate gyro 122 is typically a strap down rate gyro responsiveto the yaw axis of the aircraft. More precisely, the azimuth rate gyro122 is responsive to the azimuth rate multiplied by the cosine of thebank angle. One advantage of the present invention is that the precisebank angle does not need to be known. The feedback 120 determines thecorrect amount of azimuth rate gyro information to be subtracted frompitch rate gyro information to yield a correct steady state verticalaxis signal. The feedback also compensates for variations in thesensitivity of the azimuth rate gyro. The gain in the feedback loop maybe set by an amplifier 130.

Since the system can tolerate variation in the sensitivity of theazimuth rate gyro, the azimuth rate gyro may be an inclined rate gyro asused in a turn coordinator or turn and bank. An inclined angle of 20 to40 degrees may be used in a turn and bank to provide roll response inanticipation of azimuth response. Such roll response is not adverse tothe operation of the vertical speed information in accordance with thepresent invention. In one embodiment, turn and bank information andvertical speed information may be displayed to the pilot in the sameinstrument.

As a further advantage resulting from the feedback correction, thesystem can tolerate variations in instrument panel tilt. Aircraftinstrument panels may tilt from one aircraft model to another in theapproximate range of zero to eight degrees. The tilt potentailly altersthe amount of azimuth rate signal required to correct pitch rate in aturn, but the feedback correction automatically adjusts withoutrequiring the instrument installer to make adjustments—simplifyinginstallation and improving reliability.

The multiplier 126 may be a gain controlled amplifier, a multiplier, aduty cycle controlled pulse width modulator or other multiplyingfunction. In a digital embodiment, the multiplier 126 typically includesa multiply operation. In one embodiment, the feedback is used to varythe gain of the multiplier 126 from a nominal value. For example, zerodifference may command a nominal gain value of, for example, 50% gain. Apositive difference may increase the gain, and a negative difference maydecrease the gain. In another embodiment, the gain may be limited to arange, for example zero to one.

The filter 112 is typically a lag filter. In one embodiment, the filter112 is a single pole lag filter with a time constant of 5 seconds. Othertime constants may be used. Other multipole filters may be used. Thefilter 112 acts as a complementary filter, mixing short term pitch rategyro information 110 with long term vertical speed information 104. Thefilter 112 also integrates pitch rate gyro information 110 to generatepitch attitude angle information. The pitch angle information thendecays according to the time constant of the filter. Thus, the filteredoutput 114 is responsive to the pitch angle information, decaying at thetime constant of the filter 112 and is responsive to the vertical speedinformation 104 lagged by the time constant of the filter 112. Thus, theshort term response is primarily to the gyro based information 110 andthe long term response is primarily to the vertical speed information104. The operation of the filter may be better understood with referenceto FIG. 2.

FIG. 2A–FIG. 2E illustrate the filter response to various exemplaryvertical speed and pitch rate gyro signals. Each of the waveforms inFIGS. 2A–2E are shown using a common time axis 204 for comparison of thesignals.

FIG. 2A illustrates an exemplary output 110, 202 of the pitch rate gyro108 for a pitch rotation of one degree per second for one second. FIG.2B illustrates an exemplary vertical speed output 104, 206 resultingfrom the pitch rotation of FIG. 2A. The vertical speed of the aircraftis initially delayed relative to the control input because the aircraftmass must be accelerated to the new vertical speed. Also, vertical speedsensors may include additional lag in the sensing and display ofvertical speed and in the filtering of noise in the vertical speedsignal. Because of the lag in indicated vertical speed 206, flying thepitch axis of an airplane on vertical speed alone is very difficult. Apilot would tend to over control, resulting in severe hunting oroscillating maneuvers. At best, if the pilot can avoid over controlling,extreme concentration is required, reducing attention to other flightinstruments and increasing pilot fatigue.

FIG. 2C illustrates the output of the filter 114, 208 in response to thepitch rate gyro signal 202 of FIG. 2A. Note the initial ramp as thepitch rate gyro signal is integrated, resulting in a pitch angle signal.The pitch angle signal decays in accordance with the decay rate of thefilter.

FIG. 2D illustrates the filter output 114, 210 resulting from thevertical speed signal 206 of FIG. 2B. The filter output 210 is thevertical speed signal of FIG. 2B lagged in accordance with the timeconstant of the filter.

FIG. 2E illustrates the filter response to an exemplary maneuverincluding the response to the pitch rate gyro signal 202 of FIG. 2A andthe resulting vertical speed signal 206 of FIG. 2B. The signal 212 shownin FIG. 2E is the sum of FIGS. 2C and 2D.

FIG. 2E initially follows the ramp of FIG. 2C showing the change inpitch attitude. As the output response 208 to pitch attitude decays, theresponse 210 to vertical speed increases, maintaining a substantiallyconstant response output 212. Note that the sensitivity to pitch angle208 is adjusted to indicate the expected vertical speed for the givenpitch angle.

The resulting signal 114, 212 is a gyro enhanced vertical speed signal114, absent the lag associated with the pure vertical speed signal 206.The gyro enhanced vertical speed signal 114 may be used by a pilot muchas a pilot would use an attitude signal from a vertical gyro, resultingin stable flight, free from hunting and oscillation. Gyro enhancedvertical speed 114 has the additional advantage that the display 116indicates the vertical speed associated with the current aircraftattitude, simplifying the task of observing and coordinating multipleinstruments to fly a desired vertical speed.

FIG. 3 shows one embodiment of a vertical flight information sensor asused by the present invention. In the embodiment shown in FIG. 3, thevertical flight information sensor 102 is a vertical speed sensor. Asshown, the vertical speed signal 104 is derived from a pressurealtimeter 302 by differentiating the pressure altitude signal in adifferentiator 304 (also called a rate network 304) to derive a verticalspeed signal 104, (also called a vertical rate signal 104.)

FIG. 4 is a schematic diagram for an exemplary analog embodiment of thepresent invention. Referring to FIG. 4, the vertical information sensor102 is a vertical speed sensor comprising a pressure altitude sensor 302and a differentiating network 304, also called a rate network. The ratenetwork comprises C1, R2 and amplifier A1. Amplifier A1 is an op ampwith the + input connected to a bias reference (not shown). Thus, C1, R2and A1 form a differentiating amplifier, with the output proportional tothe rate of change of the input. R1 and C2 serve to reduce the magnitudeof high frequency noise. In one embodiment, the time constant for C2, R2is 0.5 second. The summing node 106 is at the input to A2, summing thevertical speed signal 104, the pitch rate gyro signal 110, and thecorrection signal 124. A2 and feedback network C3, R4 form the filter112. In one embodiment, the time constant, C3, R4 is 5 seconds. Thefilter output 114 may drive a display 116 or may be used in anautopilot. A feedback error signal 120 is generated at the junction ofR5 and R6. R5 and R6 are in the same ratio as R3 and R4 to compare thevertical speed signal 104 and the filter output signal 114. Thecomparison result 120 is buffered by A3 (A3 has a finite gain, forexample: one) and used to control the duty cycle of a pulse widthmodulator (PWM) 402.

The output 123 of the azimuth rate gyro 122 is rectified by the absolutevalue circuit 124 comprising R9, R10, A4, D1 and D2. The rectified pitchrate gyro signal is modulated by the switch 404 according to the dutycycle of the PWM 402 and amplified by R11, R12 and A5 to generate thecorrection signal 124. The average value of the correction signal isproportional to the duty cycle of the PWM and proportional to theabsolute value of the azimuth rate gyro and therefore proportional tothe product of the absolute value of the azimuth rate gyro signal 123and the difference 120 between the vertical speed signal 104 and thefiltered output signal 114.

In the schematic of FIG. 4, the amplifiers A1, A2, A4 and A5 arenegative gain op amp type amplifiers with the positive input of theamplifier tied to a bias reference, such as ground. Amplifier A3 is afinite gain buffer. In practice, the circuit of FIG. 4 may include otherhigh frequency roll off networks to reduce noise.

The feedback path gain comprising A3 and the sensitivity of the PWMshould be kept low—just high enough to provide good operation. If thegain is too low, the pilot may gain or lose altitude in turns. If thegain is too high, the vertical information 104 (vertical speed signal)will dominate the output 114 prematurely, introducing lag into theoutput 114. Flight tests or simulations may be used to optimize thevalue.

FIG. 5 is an exemplary block diagram of an alternate verticalinformation source 102. The vertical information 104 resulting from thesource of FIG. 5 is pitch attitude. When the source of FIG. 5 isutilized in the system of FIG. 1, the filtered signal 114 that resultsis a gyro stabilized pitch attitude. Referring to FIG. 5, an inertialaccelerometer 502 is used to sense tilt in the pitch axis, which ispitch attitude. The inertial accelerometer 502, however, is alsosensitive to axial acceleration. An alternate axial acceleration signal507 is used to remove the axial acceleration component of the inertialaccelerometer signal 503. As shown in FIG. 5, an airspeed sensor 504,generates an airspeed signal that may be differentiated in a ratenetwork 506 to derive an axial acceleration signal 507 that is notsensitive to pitch attitude. Alternatively, GPS velocity may be used togenerate axial acceleration. Axial acceleration 507 is subtracted frominertial acceleration 503 at a summing node 508 to derive a pitchattitude signal 104. Alternative inertial acceleration sensors 502include mass and spring type accelerometers, piezoresistiveaccelerometers, pendulum based accelerometers, and fluid potentiometertype sensors.

The system of FIG. 1 utilizes rate gyro 108, 122 based pitch attitudeinformation to stabilize the inertial pitch attitude verticalinformation 104 derived from the sensor of FIG. 5. The pitch attitudesignal of FIG. 5 has good absolute knowledge of pitch attitude, but mayinclude noise due to transient response in turbulence. By using thepitch attitude sensor of FIG. 5 in the system of FIG. 1, the filteredoutput 114 of FIG. 1 is a gyro stabilized pitch attitude signal. Thegyro stabilized pitch attitude signal 114 may be used to drive a display116 for the pilot may be used in an autopilot.

FIG. 6 illustrates a processor based embodiment of the presentinvention. Referring to FIG. 6, the vertical speed 102, pitch rate 108,and azimuth rate 122 signals are converted to digital using an A/Dsubsystem 604 and processed by a processor 602. The result 114 isdisplayed 116 or used in an autopilot (not shown). The system of FIG. 1may be implemented in the processor 604 by using sampled data filterssuch as IIR or FIR filters as are known in the art. The display 116 maybe an electromechanical display or may be a computer graphic displaysuch as an LCD, plasma display, or other computer graphic display knownin the art. Alternatively, the system may be configured and partitionedinto analog and digital subsystems at the convenience of the designer asis known in the art.

FIG. 7 illustrates an exemplary embodiment that includes velocityinformation. The magnitude of the pitch rate gyro signal is adjusted sothat the pitch attitude signal resulting from filtering (integrating)the pitch rate gyro signal is equal to the vertical velocity expected toresult from the pitch attitude signal. Vertical velocity, however, issubstantially proportional to the aircraft forward velocity. Thus, toaccommodate a wide range of forward velocity, it may be desirable toprovide a correction for variation in forward velocity. For manyaircraft, it may be sufficient to use a fixed gain value based on anominal cruise velocity for scaling the pitch rate gyro signal, however,for some high performance aircraft the addition of a correction signalmay be desirable.

Referring to FIG. 7, a velocity sensor 702 produces a velocity signal704 that is used to adjust the magnitude of the corrected pitch rategyro signal 706 by controlling the gain of gain stage 708. The gainstage 708 may include a multiplier, controllable gain amplifier,multiply operation, or other multiply function. The velocity signal maybe derived from airspeed or from GPS velocity.

Instrument

In one embodiment of the invention, the gyro enhanced vertical speedsignal 114 is used to drive a display 116. An exemplary display 116 isshown in FIG. 8. Referring to FIG. 8, the display 116 may be anelectromechanical display. The display 116 may show an airplane 802 thatmoves vertically in the foreground relative to a horizon line 804 in thebackground. Alternatively, the horizon line 804 may move verticallyrelative to a fixed airplane 802 in the foreground. Typically, theairplane 802 remains fixed in roll angle relative to the aircraft(instrument) frame. Alternatively, the display may be a computer graphicdisplay showing an airplane 802 and a horizon line 804.

The instrument showing gyro enhanced vertical speed 114 may includeother information as well. In one embodiment, the azimuth rate gyro 122is inclined and used to drive a turn and bank display which rotates thehorizon line 804. The azimuth rate gyro may also be used to for thecorrection signal 124 for the gyro enhanced vertical speed 114. Bothdisplays may be in the same instrument. The turn and bank display may beused to rotate the horizon line 804 in roll behind the moving airplane802 which shows the gyro enhanced vertical speed 114. The sameinstrument may also show heading or GPS course 806 as a digital numberor moving tape, or may show altitude as a number or moving tape (notshown.)

With gyro enhanced vertical speed 802, turn and bank 804, and headinginformation 806 in the same instrument 116, the pilot may use thecombined instrument as the primary flight control instrument, i.e., thepilot focus greatest attention on the combined instrument to maintainlevel flight, enter and maintain turns, terminate turns, ascend ordescend to a desired altitude. If altitude is displayed separately, onlyoccasional glances to the altimeter are necessary.

FIG. 8 illustrates an exemplary embodiment of an attitude directionindicator (ADI) instrument in accordance with the present invention.Referring to FIG. 8, the instrument is showing a climbing right turn ata heading of 120 degrees. Gyro enhanced vertical speed is indicated byan airplane FIG. 802 that moves vertically in the display in response tothe gyro enhanced vertical speed signal 114, while remaining fixedrelative to the frame in roll. Azimuth rate is indicated by a rotationof the horizon line 804 proportional to the azimuth rate signal 123. Therotation of the horizon line is suggestive of roll angle, but notstrictly indicative of roll angle. The azimuth rate gyro 122 used forturn and bank is typically inclined at a 20 to 40 degree angle to show aturn upon beginning the roll into the turn. A heading angle is indicatednumerically 802. In one embodiment, the heading angle is derived fromGPS course information. The instrument displays sufficient informationto maintain stable pitch and roll axes and to fly a given coursedirection.

Autopilot

In one embodiment of the invention, the gyro enhanced vertical speedsignal 114 is incorporated into an autopilot. When used in an autopilot,the gyro enhanced vertical speed signal 114 may also drive an optionaldisplay 116, if so configured.

In one autopilot embodiment, a vertical speed command signal is summedwith the vertical speed sensor signal 104 to produce a vertical speederror signal. The vertical speed error signal is used in place of thevertical speed sensor signal 104 of FIG. 1. The resulting gyro enhancedvertical speed signal 114 is used to command the pitch axis servo systemin the aircraft.

In a second autopilot embodiment, an altitude error signal is summedwith the vertical speed sensor signal 104 to produce a vertical speederror signal. The vertical speed error signal is used in place of thevertical speed sensor signal 104 of FIG. 1. The resulting gyro enhancedvertical speed signal 114 is used to command the pitch axis servo systemin the aircraft. The altitude error signal may be generated as thedifference between a pilot selected altitude and the measured altitude.In one embodiment the pilot may dial in a selected altitude using aselector. In another embodiment, the pilot may select the presentaltitude by pressing a button that stores the current altitude as theselected altitude.

VARIATIONS

The preferred embodiments may be varied by one skilled in the art basedon the teachings herein. Such variations may include, but are notlimited to such operations as moving functions through summing nodes,such as filtering before summing with multiple filters vs. filteringafter summing with a single filter.

In one embodiment, the gain control is a nominal value for zerodifference signal and varied around the nominal value by the differencesignal. The nominal coupling is equivalent to a fixed coupling from theabsolute value to the summing junction with a readjustment of thenominal value of the gain control.

Pitch attitude information may be derived from a separate integration ofa pitch rate gyro and the result combined with vertical information byusing complementary filter techniques. Alternatively, pitch attitudefrom a vertical gyro may be combined with vertical information usingcomplementary filter techniques.

The display may be electromechanical, computer graphic, numeric, orother display as may be devised for pilots. The symbolic representationmay be an airplane relative to a horizon line or other icon or symbol asmay be devised to communicate the information to a pilot.

The system may include digital or analog components or a mixture ofdigital and analog components and digital processing may be substitutedfor analog processing as is known by one skilled in the art.

CONCLUSION

While particular embodiments of the invention have been described, itwill be understood, however, that the invention is not limited thereto,since modifications may be made by those skilled in the art,particularly in light of the foregoing teachings. It is, thereforecontemplated by the appended claims to cover any such modifications thatincorporate those features or those improvements which embody the spiritand scope of the present invention.

1. A method for generating a gyro enhanced vertical speed signal forcontrol of an aircraft, said method comprising the steps of: combining avertical speed signal, a pitch rate gyro signal, and a correction signalto generate a combination signal; filtering said combination signal togenerate said gyro enhanced vertical speed signal; wherein saidcorrection signal is responsive to the absolute value of an azimuth rategyro signal scaled in accordance with the difference between thevertical speed signal and the gyro enhanced vertical speed signal. 2.The method of claim 1, further including a forward velocity signalwherein said absolute value of an azimuth rate gyro signal is furtherscaled based on said forward velocity signal.
 3. The method of claim 1,further including a forward velocity signal wherein said pitch rate gyrosignal is scaled based on said forward velocity signal.
 4. A system forgenerating a gyro enhanced vertical speed signal, said systemcomprising: a vertical speed sensor producing a vertical speed signal; apitch rate gyro producing a pitch rate signal; an azimuth rate gyroproducing an azimuth rate signal; a filter coupled to the vertical speedsensor and the pitch rate gyro; said filter producing said gyro enhancedvertical speed signal; said filter further coupled to a correctionsignal; said correction signal responsive to the absolute value of theazimuth rate signal scaled according to the difference between thevertical speed signal and the gyro enhanced vertical speed signal.
 5. Asystem as in claim 1, further including a display; wherein said displayis responsive to said gyro enhanced vertical speed signal.
 6. A systemas in claim 1, wherein the azimuth rate gyro is inclined.
 7. A system asin claim 1, further including a vertical speed command signal and anautopilot wherein the vertical speed command signal is summed with thevertical speed sensor signal and the gyro enhanced vertical speed sensoris used to command the autopilot.
 8. A system as in claim 1, furtherincluding an altitude error signal and an autopilot; wherein thealtitude error signal is summed with the vertical speed sensor signaland the gyro enhanced vertical speed sensor is used to command theautopilot.
 9. A system as in claim 1, wherein the vertical speed sensorincludes an altimeter and a rate network.
 10. A system as in claim 1,wherein the scaling is performed by a pulse width modulator.
 11. Thesystem of claim 1, wherein the filter is a lag filter.
 12. The system ofclaim 1, wherein the scaling is over a range from zero to one.
 13. Amethod for generating a gyro enhanced vertical information signal forcontrol of an aircraft, said method comprising the steps of: generatinga vertical information signal; generating a first pitch attitude signalbased on a pitch rate gyro signal; generating an azimuth rate signal;generating a correction signal based on the azimuth rate signal scaledbased on the difference between the vertical information signal and thegyro enhanced vertical information signal; said correction signal forcorrecting the pitch rate gyro signal; generating a gyro enhancedvertical information signal; wherein the gyro enhanced verticalinformation signal is responsive primarily to the vertical informationsignal in the long term and responsive primarily to the first pitchattitude signal in the short term.
 14. The method of claim 13, whereinthe vertical information signal is a vertical speed signal.
 15. Themethod of claim 13, wherein the vertical information signal is a secondpitch attitude signal.
 16. The method of claim 15, wherein the secondpitch attitude signal is derived from an inertial accelerometercorrected for axial acceleration.